The development of microsatellites and nanosatellites low earth orbits requires the collection of sufficient power for onboard instruments with low weight in a low volume spacecraft. Power generation methods for very small satellites of less that ten kilograms are desirable for these small satellites. Thin film solar arrays are useful power sources for small satellites. One problem faced by these low weight and low volume spacecraft is the collection of sufficient power for onboard instruments and propulsion. Body-mounted solar cells may be incapable of providing enough power when the overall surface area of a microsatellite or nanosatellite is small. Deployment of traditional planar rigid large solar arrays necessitates larger satellite volumes and weights and also requires extra apparatus needed for attitude pointing. One way to provide power to a small spacecraft is the use of roughly spherical deployable power system such as a solar powersphere that offers a relatively high collection area with low weight and low stowage volume without the need for a solar array pointing mechanism. The powersphere deployment scheme requires a deployment hinge that would move the individual hexagon and pentagon flat panels of the powersphere from a stacked configuration to an unfolded configuration where the individual panels would form a spherical structure resembling a soccer ball upon completion of the deployment sequence. The powersphere requires deployment hinges that serve to move the individual hexagon and pentagon flat panels of the powersphere from a stacked configuration to an unfolded configuration where the individual panels would form a spherical structure upon completion of the deployment sequence. Each of the panels has at least one hinge to adjacent panels. The panels should be locked into place and maintained at a precise angle relative to each connected panel to form the spherical shape. The flat hexagon and pentagon panels approximate an omnidirectional sphere. A combination of hexagon and pentagon shaped panels are used to form a soccer ball panel configuration when fully deployed. The interconnecting deployment hinges serve to position the individual flat panels of the powersphere from a stacked configuration to the deployed position forming the sphere of solar panels. The panels are hinged to one another and deploy to a precise angular position into the final shape that is preferably spherical rather than oblong or some other undesirable shape. Ideally this deployment mechanism would be fabricated from a thin film material that would have the properties to effect the mechanical positioning deployment and serve as structural elements for holding and locking each of the panels in respective positions about the powersphere.
Another type of microsatellite having an power enclosure uses a powerbox that is a three-dimensional solar array shape having rectangular shaped flat panels that would also deploy from a stowed flat configuration into a box shape configuration. The powerbox consists ideally of similarly shaped panels interconnected with hinges fabricated from a thin material that would have the properties to perfect the mechanical deployment and also be a structural element for locking each of the panels into respective positions. Hence, the powerbox would also require hinges that serve to move and lock the flat solar panels into position during deployment. Regardless of the final exterior shape of the three-dimensional power enclosure of a nanosatellite or microsatellite, a hinge mechanism is needed for deployment of the flat solar panels to cause the transition from the stowed configuration to the desired final array shape. Hence, there exists a need for positioning hinges between the flat panels forming a power collecting enclosure formed from the deployed solar array flat panels to realize any number of complex three dimensional solar array exterior surfaces used for solar power collection. However, the interconnecting hinges present a power conduction problem of routing collected converted power from the flat solar array panels to the payload of the spacecraft. Electrical conductivity of the hinge could be used to route signals and power about the power enclosure without the use of separate power lines for communicating power from the solar cell panels to the spacecraft payload. The hinges should be made of conventional materials. The hinge material could be a polymer as a flexure type hinge. But polymers are unstable and relax by cold flowing when stressed for any length of time. Polymer materials can also have undesirable outgassing properties and are generally not good electrical conductors. Polymer materials also have very low Young's moduli that reduces the deployment energy that can be stored in the hinge while stowed and later used to deploy and position the panels. Spring metals such as hardened stainless steel, beryllium copper or phosphorous bronze are commonly used as flexure type spring hinges. These spring metals have large Young's moduli, low outgassing characteristics, good electrical conductivity and will not cold flow, but spring metals have very small maximum elastic strains of 1% or less, and hence are unsuitable as deployment hinges because the steel spring hinges with interconnected panels will not stow compactly. These and other disadvantages are solved or reduced using the invention.